Contra rotatable turbine system

ABSTRACT

A three shaft ducted fan gas turbine engine ( 10 ) has a turbine system comprising a high pressure turbine rotor ( 32 ), an intermediate pressure turbine rotor ( 36 ) and first and second low pressure turbine rotors ( 42,44 ), that straddle intermediate pressure turbine rotor ( 36 ). Rotors ( 42,44 ) are mechanically connected by a casing ( 46 ) that bridges turbine rotor ( 36 ) so as to achieve transmission of torque from the first to the second thereof. Rotor ( 42 ) also delivers gas flow to the intermediate pressure rotor ( 36 ) at a much lower temperature than is experienced by conventional arrangements.

The present invention relates to turbine systems of the kind havingefficacy in aircraft powered by ducted fan gas turbine engines.

Present ducted fan gas turbine engine turbine systems compriserespective high pressure, intermediate pressure and low pressure turbinestages arranged in flow series and spaced from each other by respectivestages of fixed stators. The turbine stages rotate one each of two ormore shafts in a common direction, respective shafts being fixed tocompressors or a ducted fan.

There is documentary evidence e.g. G.B. patent 2,129,502 and U.S. Pat.No. 5,274,999, that consideration has been given to multi shaft, multistage turbine systems, in which a fan is driven by low pressure stagesthat are, in all cases, downstream of the core gas turbine system, androtate in a direction contrary to the remainder of the turbine stages.However, none of the turbine systems devised have proved sufficientlyadvantageous to warrant their manufacture and use.

The present invention seeks to provide an improved ducted fan gasturbine engine multi stage turbine system, low pressure stages of whichrotates in a direction contrary to the remainder thereof.

According to the present invention a ducted fan gas turbine engineturbine system comprises, in flow series, a high pressure bladed rotor,a fixed, bladed stator stage, a first low pressure bladed rotor, anintermediate pressure bladed rotor and a second low pressure bladedrotor, the two low pressure rotors being mechanically connected forco-rotation in a direction contrary to that of the high pressure andintermediate pressure rotors, said contra rotation being effected byappropriate shaping of the gas flow paths through said stage of statorblades.

The invention will now be described, by way of example and withreference to the accompanying drawings, in which:

FIG. 1 is an axial cross section through a ducted fan gas turbine engineincluding a turbine system in accordance with one aspect of the presentinvention.

FIG. 2 is an axial cross section through a ducted fan gas turbine enginencluding a turbine system in accordance with a further aspect of thepresent invention.

FIG. 3 is an axial cross section through a ducted fan gas turbine engineincluding a turbine system having an alternative fan drive system inaccordance with the present invention.

Referring to FIG. 1. A ducted fan gas turbine engine, generallyindicated by the numeral 10, has a short fan cowl 12 that surrounds astage of fan blades 14. A casing 16 supports the fan cowl 12 via struts18, and also surrounds the core gas generator 20.

The core gas generator 20 consists of the following: An intermediatepressure compressor 22, that via fixed vanes 24, receives and furtherpressurises ambient air after it has been slightly pressurised onpassing between the root portions 26 of the rotating fan stage 14. Therefollows a high pressure compressor 28 that further compresses the air onits receipt from the intermediate compressor 22, and delivers some of itat a still higher pressure to combustion equipment 30 wherein it ismixed with fuel and burned. A high pressure turbine roter 32 receivesthe resulting hot gases and is caused to rotate in a given directionand, being directly connected to high pressure compressor 28 via a shaft34, rotates it in the same direction. The final part of the core gasgenerator 20 is an intermediate pressure turbine rotor 36 whicheventually receives the gases from the high pressure rotor 32 at apressure lower than that experienced by high pressure rotor 32. The gasflow path is such that the intermediate pressure turbine rotor 36rotates in the same direction as the high pressure turbine rotor 32.Intermediate pressure turbine rotor 36 is directly connected tointermediate pressure compressor 22 via a shaft 38 and consequentlyrotates intermediate compressor 22 in the same direction.

Fan stage 14 is not mechanically connected to the core gas generator inany way for the purpose of rotation thereby, but derives its rotarymotion via a shaft 40 that is directly connected to the downstream one,44, of two low pressure power turbine rotors 42 and 44. By “downstream”is meant with regard to the direction of flow of gases through theengine. Turbine rotors 42 and 44 are mechanically connected via an outercasing 46 that co-rotates with them.

Turbine rotors 42 and 44 straddle intermediate turbine rotor 36, rotor42 being positioned upstream thereof, and consequently is the firstturbine rotor to receive the flow of gas that has passed through highpressure turbine 32, and thereafter through a stage of fixed, wide chordstators 48. The stators 48 are shaped so as to define flow paths betweenthem which will increase the exit whirl of the gases leaving the stators48 and thereby reduce the angle of deflection of the gases relative tothat imposed by conventional stators, and thus makes for a moreaerodynamically efficient stator stage.

The low pressure turbine rotor 42 is arranged so that on receiving theflow from stator stage 36, it counter rotates relative to the highpressure turbine rotor 32. However, the intermediate pressure turbinerotor 36 is arranged so that, on receiving the flow from low pressureturbine rotor 42, it rotates in the same direction as the high pressureturbine rotor 32. The second low pressure turbine rotor 44, is rotatedby torque transmitted by low pressure turbine stage 42, via casing 46and, to some extent, by the gases exiting the intermediate turbine stage36.

During the cruise regime of an aircraft powered by a known three shaftducted fan engine, in which all the turbines and their respective shaftsrotate in a common direction, the ratio of the speeds of the highpressure turbine shaft, the intermediate pressure turbine shaft and thelow pressure turbine shaft is, nominally, 11:7:3. Thus, the speed of thelow pressure shaft is less than half that of the intermediate shaftspeed. However, in the three shaft arrangement of the present invention,whilst the ratio of the speeds of the high and intermediate pressureshafts will nominally remain the same as the prior art engine, the speedratio between the intermediate pressure shaft and low pressure shaftwill increase to, nominally, 10, by virtue of their effected counterrotation. This may necessitate the substitution of prior art qualityinter shaft bearings (not shown) between intermediate pressure shaft 38and low pressure shaft 40, by bearings (not shown) of the same qualityas those used to support high pressure shaft 34 relative to surroundingfixed structure in known manner.

During operation of an associated engine, the contra rotation ofintermediate pressure turbine rotor 36 and low pressure turbine rotor 42results in the speed of rotation of the blades of intermediate pressurerotor 36 being much higher than that of the blades of low pressure rotor42. A very high exit whirl velocity of the gases leaving the blades oflow pressure rotor 42 is achieved, and still maintain good inletconditions into intermediate pressure rotor 36. By this means, a largefraction, in the order of 40-50% of the total power generated by the lowpressure system of turbine rotors 42 and 44, can be generated by turbinerotor 42.

The FIG. 1 example of the present invention depicts a further,conventional low pressure turbine rotor 50 with stators 52. Theinclusion of this or any other number of conventional low pressureturbine rotors will depend on the required power regime of the ductedfan gas turbine engine, and the example should not be regarded aslimitive.

A number of advantages may be derived from the turbine system of thepresent invention as follows:

-   -   a) No stator stages are utilised either side of the intermediate        pressure turbine rotor 36, thus achieving a large cost saving.    -   b) The stator inlet to the low pressure turbine rotor 42 becomes        a low deflection aerofoil, thereby improving its aerodynamic        efficiency.    -   c) Low pressure turbine rotor 42 and intermediate pressure        turbine rotor 36 both have velocity ratios that are much higher        than conventional turbines, thus helping to achieve high        aerodynamic efficiency.    -   d) The respective frequencies of the aerodynamic interaction        between the intermediate pressure turbine rotor and the low        pressure turbine rotors 42 and 44 are much higher than in        conventional designs. This helps to keep noise generated by the        interactions out of the audible range, and reduces the perceived        noise of the turbine system.    -   e) A substantial reduction in the number of low pressure turbine        blades is achieved where the values of aerofoil coefficients and        axial chord proportions are similar to those used in a        conventional turbine design, e.g. the reduction could equal        about 30% relative to the number of blades utilised in a five        stage conventional low pressure turbine.    -   f) Flowing the hot gases through stators 48 and low pressure        turbine rotor 42, prior to it reaching intermediate pressure        turbine rotor 36, results in the gas experiencing a considerable        drop in temperature. The intermediate pressure turbine rotor 36        is subjected to considerably lower temperatures than in        conventional designs. This invention does not reduce its stress.        Moreover, higher stress levels can be allowed in the IP rotor.        This is useful because in the current invention the flow will        have expanded more, relative to conventional, by the time it        leaves the IP rotor. To achieve an optimum design, the flow area        of the IP rotor needs to be larger than conventional, which        results in higher stress.

The first low pressure rotor 42 will experience a higher temperaturethan in a conventional design. However, since it is on the LP shaftwhich rotates at the lowest speed, the levels of stress are low anddesigns can be achieved using existing materials without the need forcooling.

Referring to FIG. 2. In this second embodiment of the present invention,the turbine system is generally as that described with reference toFIG. 1. However, the intermediate pressure shaft 38 is provided with agear 54 about its periphery. Gear 54 is engaged by planet gears 56 thatare supported by fixed structure 57 passing through stators 48. Lowpressure turbine rotor 42 is provided with a gear ring 58 that alsoengages planet gears 56. The ratio of gears can be determined usingknown gear design technology.

The utilising of a set of gears 54, 56 and 58 enables the intermediatepressure turbine to extract from the gas flow, the power to drive theintermediate compressor 22, and the power to drive the low pressureturbine rotor 42 through the gears. There results an increase in thewhirl velocity in the gas flow leaving the intermediate pressure turbinerotor, thereby improving the inlet conditions into the low pressureturbine rotor 44, which is then able to extract more power from the gasflow than is possible in the first embodiment described herein, whilestill achieving acceptable aerodynamic efficiency.

The FIG. 2 embodiment of the present invention will require fewer lowpressure turbine rotors than the embodiment of FIG. 1. Further, if thevalues of blade lift coefficients and axial chord proportions aresimilar to those in a conventional low pressure turbine rotor design, aneven greater reduction of approximately 60% in the number of lowpressure turbine blades can be achieved. Moreover, the number of bladesin the intermediate pressure turbine rotor 36 can be reduced bytypically 10%.

The operation of the FIG. 3 embodiment in a relatively cool part of theengine enables cooling of the gear system to be more easily effected.

In this embodiment, the geardrive is located at the front of the engine,between the fan 14 and the intermediate compressor 22, and the fan discis provided with a gear ring 158. Gear ring 158 is engaged by planetgears 156 that are supported by fixed structure 157 passing through thefixed vanes 24 at inlet to the IP compressor 22. The IP compressor 22 isprovided with another gear ring 154 that also engages planet gears 156.As before, the ratio of gears can be determined using known gear designtechnology.

1. A ducted fan gas turbine engine turbine system comprising, in flowseries, a high pressure bladed turbine rotor, a fixed, bladed statorstage, a first low pressure bladed turbine rotor, an intermediatepressure bladed turbine rotor and a second low pressure bladed turbinerotor, the two low pressure bladed turbine rotors being mechanicallyconnected for co-rotation in a direction contrary to that of the highpressure and intermediate pressure turbine rotors, said contra rotationbeing effected, at least in part, by appropriate shaping of the gas flowpaths through said stage of stator blades.
 2. A ducted fan gas turbineengine turbine system as claimed in claim 1 wherein said mechanicalconnection comprises a casing fixed on the outer tip extremities of theblades of each low pressure turbine rotor and bridging said intermediatepressure turbine rotor.
 3. A ducted fan gas turbine engine turbinesystem as claimed in claim 1 wherein in situ in a ducted fan gas turbineengine, said high pressure turbine rotor is drivingly connected via ashaft to a high pressure compressor, said intermediate pressure turbinerotor is drivingly connected via another shaft to an intermediatepressure compressor immediately upstream of said high pressurecompressor and said second low pressure turbine rotor is directlydrivingly connected via a further shaft to a ducted fan at the upstreamend of said engine.
 4. A ducted fan gas turbine engine turbine system asclaimed in any of claims 1 wherein said intermediate pressure turbinerotor is effectively mechanically connected to said first low pressureturbine rotor so as to impart a rotary force thereon during operation ina said engine.
 5. A ducted fan gas turbine engine turbine system asclaimed in claim 4 wherein said mechanical connection comprises a gearedconnection between the intermediate turbine rotor shaft and said firstlow pressure turbine rotor.
 6. A ducted fan gas turbine engine turbinesystem as claimed in claim 5 wherein said geared connection comprises aspur gear on said intermediate turbine rotor shaft that engages aplurality of planet gears supported from fixed structure and which inturn engage a spur gear on said first low pressure turbine rotor.
 7. Aducted fan gas turbine engine turbine system as claimed in claim 4wherein said mechanical connection comprises a geared connection betweenthe upstream end of an intermediate compressor and a ducted fan when allare assembled.
 8. A ducted fan gas turbine engine turbine system asclaimed in claim 5 wherein said connection of said gears is made viaplanet gears supported from fixed structure.